Gas turbine engines (“GTE”) have been engineered extensively to improve efficiency, thrust-to-weight ratios, and other measures of engine performance. One of the most direct manners by which engine performance can be improved is through increases in the core rotational speeds and turbine inlet temperatures generated during engine operation. However, as turbine inlet temperatures and rotational speeds increase, so too do the thermal and mechanical demands placed on the GTE components. The most demanding performance requirements are typically placed on the high pressure turbine rotor or rotors, which are positioned immediately downstream of the GTE combustion section and which rotate at the greatest speed during engine operation. The turbine blades, in particular, are directly exposed to combustive gas flow at or near peak temperatures and are consequently heated to exceedingly high temperatures at which most alloys weaken or melt and become prone to oxidation or other forms of chemical degradation. By comparison, the inner portion of the turbine (commonly referred to as the “turbine disk”) is largely shielded from direct exposure to combustive gas flow, but is subject to considerable mechanical stress resulting from the centrifugal forces acting on the turbine rotor at high rotational speeds.
Turbines can be broadly divided into two categories, axial and radial turbines, based upon the direction of airflow received by the turbine relative to the turbine's rotational axis. Relative to axial turbines, radial turbines offer certain performance benefits including decreased thermal and mechanical stresses, which enable the turbine to operate at greater efficiencies and at higher temperatures. However, due to the nature of their design, radial turbines are relatively lengthy in the axial direction. As a result, a radial turbine can be undesirably heavy and cumbersome. Axial turbines can readily be cooled, which allows such turbines to operate at high inlet temperature, while despite the proposal of multiple cooled radial turbine concepts, there are no cooled radial turbines currently implemented in commercial gas turbine engine platforms. As a result, majority of gas-turbine engine employing radial turbine are operating at much lower turbine inlet temperature relative to those employing an axial turbine. In addition, it is often difficult to fabricate radial turbines utilizing conventional casting processes, especially if such radial turbines include relatively complex internal cooling flow passages. Still further, present blade cooling schemes are unable to adequately cool the blade tip and the “saddle” region (between blades) of the turbine, resulting in undesirable high component metal operating temperatures and temperature gradients, especially during start-up conditions. For at least these reasons, and despite the proposal of multiple radial turbine designs in the prior art, few currently-implemented gas turbine engine platforms incorporate radial turbines.
It would thus be desirable to provide a radial turbine suitable for usage in a gas turbine engine that is relatively lightweight and cost effective to manufacture and implement, can operate at elevated turbine inlet temperature levels, and that has improved blade tip and saddle region cooling characteristics. Other desirable features and characteristics of embodiments of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying drawings and the foregoing Background.